Seal arrangement for turbine engine component

ABSTRACT

A component for a gas turbine engine according to an example of the present disclosure includes, among other things, a body including a cold side surface adjacent to a mate face. A plurality of ridges extends from the cold side surface. A seal member abuts the plurality of ridges to define a plurality of cooling passages. The seal member is configured to move between a first position and a second position relative to the plurality of ridges. Each of the plurality of cooling passages includes a first inlet defined at the first position and a second, different inlet defined at the second position. A method of sealing between adjacent components of a gas turbine engine is also disclosed.

CROSS-REFERENCE TO RELATED APPLICATION

This application is a divisional of U.S. patent application Ser. No.14/704,278 filed May 5, 2015.

BACKGROUND

This disclosure relates to impingement cooling for a component of a gasturbine engine, and more particularly to a seal arrangement having oneor more ridges for cooling augmentation.

Gas turbine engines can include a fan for propulsion air and to coolcomponents. The fan also delivers air into a core engine where it iscompressed. The compressed air is then delivered into a combustionsection, where it is mixed with fuel and ignited. The combustion gasexpands downstream over and drives turbine blades. Static vanes arepositioned adjacent to the turbine blades to control the flow of theproducts of combustion. The blades and vanes are subject to extremeheat, and thus cooling schemes are utilized for each.

Adjacent blades or vanes are distributed to define leakage gaps atadjacent mate faces. Cooling airflow is communicated through the leakagegaps to cool surfaces of the mate faces.

SUMMARY

A component for a gas turbine engine according to an example of thepresent disclosure includes a body including a cold side surfaceadjacent to a mate face. A plurality of ridges extends from the coldside surface. A seal member abuts the plurality of ridges to define aplurality of cooling passages. The seal member is configured to movebetween a first position and a second position relative to the pluralityof ridges. Each of the plurality of cooling passages includes a firstinlet defined at the first position and a second, different inletdefined at the second position.

In a further embodiment of any of the forgoing embodiments, each of theplurality of cooling passages includes an outlet adjacent to the mateface.

In a further embodiment of any of the forgoing embodiments, each of theplurality of ridges includes a first end and a second end. The outlet ofeach of the plurality of cooling passages is located at the second end,and the seal member is dimensioned such that the seal member is spaced adistance from the first end.

In a further embodiment of any of the forgoing embodiments, each of theplurality of ridges includes a first passage portion transverse to asecond passage portion, and the second passage portion is configured toextend outboard of the seal member.

In a further embodiment of any of the forgoing embodiments, at leastsome of the plurality of ridges includes a radial retention featureextending from a respective one of the second passage portion. Theradial retention feature is configured to abut a first surface of theseal member opposite from a second surface of the seal member abuttingthe first passage portion.

In a further embodiment of any of the forgoing embodiments, the coldside surface is located at a slot extending inwardly from the mate face,and the slot is configured to receive the seal member.

In a further embodiment of any of the forgoing embodiments, thecomponent is one of an airfoil, a blade outer air seal (BOAS), and acombustor panel.

In a further embodiment of any of the forgoing embodiments, the airfoilincludes an airfoil section extending from a platform, and the cold sidesurface is located at an undersurface of the platform.

In a further embodiment of any of the forgoing embodiments, the airfoilis a turbine blade.

A gas turbine engine according to an example of the present disclosureincludes a first component including a first set of ridges protrudingfrom a first cold side surface adjacent to a first mate face, and asecond component including a second set of ridges protruding from asecond cold side surface adjacent to a second mate face. The second mateface is circumferentially adjacent to the first mate face to define aleakage gap. A seal member abuts the first set of ridges to define afirst set of cooling channels, and abuts the second set of ridges todefine a second set of cooling channels. The seal member is spaced apartfrom the first cold side surface and the second cold side surface.

In a further embodiment of any of the forgoing embodiments, each of thefirst set and the second set of cooling channels includes an inletspaced apart from the leakage gap and an outlet adjacent to the leakagegap.

In a further embodiment of any of the forgoing embodiments, each of theplurality of ridges includes a first passage portion transverse to asecond passage portion, and the second passage portion is configured tobound relative movement of the seal member in a circumferentialdirection.

In a further embodiment of any of the forgoing embodiments, the sealmember defines a first width in a circumferential direction, andoutermost distal ends of the first set and the second set of ridgesdefine a second width in the circumferential direction, and a ratio ofthe first width to the second width is equal to or less than 0.8.

A further embodiment of any of the foregoing embodiments includes theseal member including a sealing surface configured to abut the first setand the second set of ridges, and an outer surface spaced apart from thesealing surface. The first component and the second component are spacedfrom the outer surface at each circumferential position of the sealmember.

In a further embodiment of any of the forgoing embodiments, each of thefirst component and the second component is one of an airfoil and ablade outer air seal (BOAS).

In a further embodiment of any of the forgoing embodiments, the firstcomponent is an airfoil. The airfoil includes an airfoil sectionextending from a platform. The platform includes an upper surfacebounding a core flow path and an undersurface bounding a cooling cavity,and the first cold side surface is located at the undersurface of theplatform.

A method of sealing between adjacent components of a gas turbine engineaccording to an example of the present disclosure includes positioning afeather seal across a leakage gap defined between mate faces of adjacentcomponents, and along a plurality of ridges to define a plurality ofcooling passages. The plurality of ridges are configured to protrudefrom cold side surfaces of the adjacent components such that the featherseal is spaced apart from the cold side surfaces.

A further embodiment of any of the foregoing embodiments includescommunicating coolant from a cooling cavity to the plurality of coolingpassages when an edge face of the feather seal is substantially alignedwith ends of at least some of the plurality of ridges.

In a further embodiment of any of the forgoing embodiments, each of theplurality of cooling passages includes an inlet that is spaced apartfrom the mate faces at each position of the feather seal relative to theplurality of ridges.

In a further embodiment of any of the forgoing embodiments, each of theplurality of ridges includes a first passage portion transverse to asecond passage portion, and the second passage portion is configured tobound relative movement of the feather seal along the plurality ofridges.

Although the different examples have the specific components shown inthe illustrations, embodiments of this disclosure are not limited tothose particular combinations. It is possible to use some of thecomponents or features from one of the examples in combination withfeatures or components from another one of the examples.

The various features and advantages of this invention will becomeapparent to those skilled in the art from the following detaileddescription of an embodiment. The drawings that accompany the detaileddescription can be briefly described as follows.

BRIEF DESCRIPTION OF THE DRAWINGS

FIG. 1 schematically shows a gas turbine engine.

FIG. 2 schematically shows an airfoil arrangement for a turbine section.

FIG. 3A illustrates a side view of a first embodiment of a coolingarrangement for an airfoil.

FIG. 3B illustrates a cross-sectional view of the cooling arrangementalong line 3B-3B of FIG. 3A.

FIG. 3C illustrates a bottom view of selected portions the coolingarrangement of FIG. 3B with a seal member in a first position.

FIG. 3D illustrates a bottom view of selected portions the coolingarrangement of FIG. 3B with the seal member in a second position.

FIG. 4A illustrates a side view of a cooling arrangement for an airfoiland having a plurality of ridges with various geometries.

FIG. 4B illustrates a side view of a cooling arrangement for an airfoilhaving a plurality of ridges with various geometries.

FIG. 5A illustrates a side view of a second embodiment of a coolingarrangement for an airfoil.

FIG. 5B illustrates a cross-sectional view of the cooling arrangementalong line 5B-5B of FIG. 5A.

FIG. 5C illustrates an isometric view of selected portions of theairfoil of FIG. 5A.

FIG. 6A illustrates a side view of a cooling arrangement for a vane.

FIG. 6B illustrates a cross-sectional view of the cooling arrangementalong line 6B-6B of FIG. 6A.

FIG. 7 illustrates a cross-sectional view of a cooling arrangement foran engine component.

FIG. 8A illustrates a side view of a cooling arrangement for a vane.

FIG. 8B illustrates a cross-sectional view of the cooling arrangementalong line 8B-8B of FIG. 8A.

FIG. 9 illustrates a cross-sectional view of a cooling arrangement foran engine component.

FIG. 9A illustrates a perspective view of selected portions of theengine component of FIG. 9.

DETAILED DESCRIPTION

FIG. 1 schematically illustrates a gas turbine engine 20. The gasturbine engine 20 is disclosed herein as a two-spool turbofan thatgenerally incorporates a fan section 22, a compressor section 24, acombustor section 26 and a turbine section 28. Alternative engines mightinclude an augmentor section (not shown) among other systems orfeatures. The fan section 22 drives air along a bypass flow path B in abypass duct defined within a nacelle 15, while the compressor section 24drives air along a core flow path C for compression and communicationinto the combustor section 26 then expansion through the turbine section28. Although depicted as a two-spool turbofan gas turbine engine in thedisclosed non-limiting embodiment, it should be understood that theconcepts described herein are not limited to use with two-spoolturbofans as the teachings may be applied to other types of turbineengines including three-spool architectures.

The exemplary engine 20 generally includes a low speed spool 30 and ahigh speed spool 32 mounted for rotation about an engine centrallongitudinal axis A relative to an engine static structure 36 viaseveral bearing systems 38. It should be understood that various bearingsystems 38 at various locations may alternatively or additionally beprovided, and the location of bearing systems 38 may be varied asappropriate to the application.

The low speed spool 30 generally includes an inner shaft 40 thatinterconnects a fan 42, a first (or low) pressure compressor 44 and asecond (or low) pressure turbine 46. The inner shaft 40 is connected tothe fan 42 through a speed change mechanism, which in exemplary gasturbine engine 20 is illustrated as a geared architecture 48 to drivethe fan 42 at a lower speed than the low speed spool 30. The high speedspool 32 includes an outer shaft 50 that interconnects a second (orhigh) pressure compressor 52 and a first (or high) pressure turbine 54.A combustor 56 is arranged in exemplary gas turbine 20 between the highpressure compressor 52 and the high pressure turbine 54. A mid-turbineframe 57 of the engine static structure 36 is arranged generally betweenthe high pressure turbine 54 and the low pressure turbine 46. Themid-turbine frame 57 further supports bearing systems 38 in the turbinesection 28. The inner shaft 40 and the outer shaft 50 are concentric androtate via bearing systems 38 about the engine central longitudinal axisA which is collinear with their longitudinal axes.

The core airflow is compressed by the low pressure compressor 44 thenthe high pressure compressor 52, mixed and burned with fuel in thecombustor 56, then expanded over the high pressure turbine 54 and lowpressure turbine 46. The mid-turbine frame 57 includes airfoils 59 whichare in the core airflow path C. The turbines 46, 54 rotationally drivethe respective low speed spool 30 and high speed spool 32 in response tothe expansion. It will be appreciated that each of the positions of thefan section 22, compressor section 24, combustor section 26, turbinesection 28, and fan drive gear system 48 may be varied. For example,gear system 48 may be located aft of combustor section 26 or even aft ofturbine section 28, and fan section 22 may be positioned forward or aftof the location of gear system 48.

The engine 20 in one example is a high-bypass geared aircraft engine. Ina further example, the engine 20 bypass ratio is greater than about six(6), with an example embodiment being greater than about ten (10), thegeared architecture 48 is an epicyclic gear train, such as a planetarygear system or other gear system, with a gear reduction ratio of greaterthan about 2.3 and the low pressure turbine 46 has a pressure ratio thatis greater than about five. In one disclosed embodiment, the engine 20bypass ratio is greater than about ten (10:1), the fan diameter issignificantly larger than that of the low pressure compressor 44, andthe low pressure turbine 46 has a pressure ratio that is greater thanabout five (5:1). Low pressure turbine 46 pressure ratio is pressuremeasured prior to inlet of low pressure turbine 46 as related to thepressure at the outlet of the low pressure turbine 46 prior to anexhaust nozzle. The geared architecture 48 may be an epicycle geartrain, such as a planetary gear system or other gear system, with a gearreduction ratio of greater than about 2.3:1. It should be understood,however, that the above parameters are only exemplary of one embodimentof a geared architecture engine and that the present invention isapplicable to other gas turbine engines including direct driveturbofans.

A significant amount of thrust is provided by the bypass flow B due tothe high bypass ratio. The fan section 22 of the engine 20 is designedfor a particular flight condition—typically cruise at about 0.8 Mach andabout 35,000 feet. The flight condition of 0.8 Mach and 35,000 ft, withthe engine at its best fuel consumption—also known as “bucket cruiseThrust Specific Fuel Consumption (‘TSFC’)”—is the industry standardparameter of lbm of fuel being burned divided by lbf of thrust theengine produces at that minimum point. “Low fan pressure ratio” is thepressure ratio across the fan blade alone, without a Fan Exit Guide Vane(“FEGV”) system. The low fan pressure ratio as disclosed hereinaccording to one non-limiting embodiment is less than about 1.45. “Lowcorrected fan tip speed” is the actual fan tip speed in ft/sec dividedby an industry standard temperature correction of [(Tram °R)/(518.7°R)]^(0.5). The “Low corrected fan tip speed” as disclosedherein according to one non-limiting embodiment is less than about 1150ft/second.

FIG. 2 shows selected portions of the turbine section 28 including arotor 60 carrying one or more airfoils 61 for rotation about the centralaxis A. In this disclosure, like reference numerals designate likeelements where appropriate and reference numerals with the addition ofone-hundred or multiples thereof designate modified elements that areunderstood to incorporate the same features and benefits of thecorresponding original elements. Each airfoil 61 includes a platform 62and an airfoil section 65 extending in a radial direction R from theplatform 62 to a tip 64. The airfoil section 65 generally extends in achordwise direction X between a leading edge 66 and a trailing edge 68.A root section 67 of the airfoil 61 is mounted to the rotor 60, forexample. It should be understood that the airfoil 61 can alternativelybe integrally formed with the rotor 60, which is sometimes referred toas an integrally bladed rotor (IBR). A blade outer air seal (BOAS) 69 isspaced radially outward from the tip 64 of the airfoil section 65. Avane 70 is positioned along the engine axis A and adjacent to theairfoil 61. The vane 70 includes an airfoil section 71 extending betweenan inner platform 72 and an outer platform 73 to define a portion of thecore flow path C. The turbine section 28 includes multiple airfoils 61,vanes 70, and BOAS 69 arranged circumferentially about the engine axisA.

The outer platform 73 of vane 70 and BOAS 69 can define one or moreouter cooling cavities 74. The platform 62 of airfoil 61 and the innerplatform 72 of vane 70 can define one or more inner cooling cavities 75.The cooling cavities 74, 75 are configured to receive cooling flow fromone or more cooling sources 76 to cool portions of the airfoil 61, BOAS69 and/or vane 70. Cooling sources 76 can include bleed air from anupstream stage of the compressor section 24 (shown in FIG. 1), bypassair, or a secondary cooling system aboard the aircraft, for example.Each of the cooling cavities 74, 75 can extend in a thickness directionT between adjacent airfoils 61, BOAS 69 and/or vanes 70, for example.

FIGS. 3A to 3D illustrate an exemplary cooling arrangement 178 for anairfoil 161. Although the exemplary cooling arrangements discussedherein primarily refer to a turbine blade, the teachings herein can alsobe utilized for another portion of the engine 20, such as vane 70, anupstream stage of the compressor section 24, or combustor panels locatedin the combustor section 26 and defining portions of a combustionchamber, exhaust nozzles, or augmentors, for example. The exemplarycooling arrangements discussed herein can also be utilized adjacent toeither of the cooling cavities 74, 75 and at various positions relativeto the core flow path C.

Adjacent airfoils 161A and 161B have mate faces 180A, 180B arrangedcircumferentially about the engine axis A to define a leakage gap 181.The leakage gap 181 is configured to receive pressurized cooling airflowfrom cooling cavity 175 for providing cooling to the hot side surfacesof the mate faces 180A, 180B and reduce ingestion of hot gases from thecore flow path C into the cooling cavity 175. The relatively warmercooling airflow is discharged from the leakage gap 181 to the core flowpath C.

A seal member 182 including one or more segments can be arrangedadjacent to cold side surfaces 184A, 184B of the airfoil 161A, 161B, forexample, and about the engine axis A to separate the cooling cavity 175from the core flow path C. The cold side surfaces 184B, 184B of theairfoils 161A, 161B are adjacent to the mate faces 180A, 180B. In theillustrative embodiment, the cold side surfaces 184B, 184B are locatedat, or are otherwise defined by, undersurfaces 183A, 183B of platforms162A, 162B, and the seal member 182 is a feather seal arranged adjacentto undersurfaces 183A, 183B. The feather seal can be fabricated fromsheet metal made of nickel or cobalt, for example. Other materials forthe seal member 182 can be utilized, including various high temperatureNi, Cobalt, or Inco alloys, or composite materials, for example.

Each airfoil 161A, 161B includes a plurality of ridges 185A, 185Bextending radially or otherwise protruding from the cold side surfaces184B, 184B of the platforms 162A, 162B. The ridges 185 are distributedboth axially and circumferentially adjacent to the leakage gap 181 todefine a plurality of grooves 186. Each of the ridges 185 includes aproximal (or first) end 190 adjacent to the leakage gap 181 or mate face180, and a distal (or second) end 191 spaced apart from the leakage gap181 or mate face 180. In the illustrative embodiment, the ridges 185A,185B extend circumferentially from edges of the mate faces 180A, 180B.In an alternative embodiment, at least some of the ridges 185A, 185B areoffset away from edges of the mate faces 180A, 180B.

The ridges 185A, 185B can be arranged at various orientations relativeto the cold side surfaces 184A, 184B and mate faces 180A, 180B, such assubstantially perpendicular or transverse relative orientations, andpitched (spaced) axially and/or circumferentially to provide a desiredheat transfer and cooling augmentation to portions of the platform 162adjacent the cooling passages 187. The relative orientation of theridges 184A, 184B may be altered to increase wetted surface area, and todirect and regulate leakage flow to high heat load locations along theedges of the mate faces 180A, 180B. Additionally, the ridges 185A, 185Balso improve platform creep capability by providing additionalstiffening and lower local and bulk average bending stress of theplatforms 162A, 162B.

The ridges 185A, 185B can be configured having various geometries, suchas a rectangular cross-sectional profile as shown by ridges 185A, 185Bin FIG. 3A. Other geometries can include various curvatures asillustrated by ridges 285 in FIG. 4A, a trapezoidal cross-sectionalgeometry as shown by ridges 385A in FIG. 4B, fully radiused radiallyextending cross-sectional geometries as shown by ridges 385B (FIG. 4B)to minimize local stress concentrations associated with sharp edges,corners and inflection points, or a combination of geometries depictedby ridges 385A, 385B, for example. Various techniques for forming theridges 185 can be utilized. In some embodiments, the ridges 185 are castor additively manufactured. In another embodiment, the ridges 185 orgrooves 186 are machined from a portion of the cold side surface 184.

The seal member 182 is arranged to abut or span across at least aportion of the ridges 185A, 185B to define a plurality of coolingpassages 187. The seal member 182 includes one or more edge faces suchas side faces 195 at ends of the seal member 182. Other edge faces caninclude a sealing surface 196 and an outer surface 197 extending betweenthe side faces 195. The sealing surface 196 is configured to abut radialsurfaces of the ridges 185A, 185B, and the outer surface 197 is spacedapart from the sealing surface 196. In some embodiments, ridges 185A,185B are configured such that the outer surface 197 of the seal member182 is spaced apart from the cold side surfaces 184B, 184B and otherportions of the airfoil 161 at each position of the seal member 182.

Each of the cooling passages 187 includes an inlet 188 (shown in dashedline in FIG. 3B) in communication with the cooling cavity 175 adjacentto the distal end 191 of adjacent ridges 185, and an outlet 189 adjacentto the proximal end 190 of adjacent ridges 185 and in communication withthe leakage gap 181. The inlets 188A, 188B are spaced apart from themate faces 180A, 180B and leakage gap 181, and extend generally in theradial direction R, for example, between the cold side surfaces 184B,184B of the platforms 162A, 162B and the seal member 182.

The seal member 182 can be dimensioned such that the seal member 182 isspaced a width D1 in the circumferential or thickness direction T fromdistal (or first) ends 191A, 191B of the ridges 185A, 185B, as shown inFIG. 3C. In some embodiments, the seal member 182 is dimensioned todefine a width D2 in the circumferential or thickness direction T, andthe outermost distal ends 191A, 191B of the ridges 185A, 185B define awidth D3 in the circumferential or thickness direction T such that aratio of the first width to the second width is equal to or less than0.8, or more narrowly between 0.5 and 0.75. These arrangements permitadditional cooling augmentation to portions of the platform 162 adjacentthe cooling passages 187.

During operation, the seal member 182 may move circumferentiallyrelative to the ridges 185A, 185B, as illustrated by different positionsof the seal member 182 in FIGS. 3C and 3D. The seal member 182 isconfigured to move between a first position and a second positionrelative to the ridges 185A, 185B. For example, the seal member 182 canmove between a first position spaced from the mate faces 180A, 180B(FIG. 3C) and a second position adjacent to ends of the plurality ofridges 185B (FIG. 3D) or ridges 185A, or even circumferentially past theends of the ridges 185B. Accordingly, a location of each inlet 188 canchange due to relative movement of the seal member 182 along the ridges185A, 185B such that each of the cooling passages 187A, 187B includes afirst inlet 188A, 188B defined at the first position (FIG. 3C) and asecond, different inlet 188A′, 188B′ defined at the second position(FIG. 3D). In some situations, the inlets 188 are selectively defined atdistal ends 191 of the ridges 185 and at an edge face of the seal member182 such as one of the side faces 195, as illustrated by the arrangementof the seal member 182 relative to the distal ends 191B in FIG. 3C, andin other positions of the seal member 182 are spaced from the distalends 191 of the ridges 185 as illustrated in FIG. 3D. The arrangement ofthe ridges 185A, 185B spacing the cold side surfaces 184B, 184B of theplatforms 162A, 162B from the seal member 182 permits the inlets 188A,188B to have a radial component, thereby allowing cooling airflow toenter the cooling passages 187A, 187B through the inlets 188A, 188B evenwhen one or more edge faces of the seal member 182 is substantiallyaligned with the ends of at least some of the ridges 185. Thus, blockageof the inlets 188 caused by interaction of the seal member 182 and thecold side surfaces 184A, 184B can be reduced for each position of theseal member 182 relative to the ridges 185A, 185B.

Although the various cooling arrangements have been primarily discussedwith respect to airfoils or turbine blades, the teachings herein can beutilized for other portions of the engine 20, such as one or more vanes570 shown in FIGS. 6A and 6B, or one or more BOAS 669 or combustorpanels 663 shown in FIG. 7. Although FIGS. 6A and 6B illustrate acooling arrangement 578 adjacent to an outer platform 572 of vane 570,the cooling arrangement 578 can be utilized for an inner platform of avane, such as inner platform 72 of vane 70 shown in FIG. 2.

FIGS. 5A to 5C illustrate a second embodiment of a cooling arrangement478 for an airfoil 461, such as the one or more airfoils 61 of FIG. 2.The cooling arrangement 478 can also be utilized for a vane or BOAS,such as vane 70 or BOAS 69 of FIG. 2. In the illustrative embodiment,one or more ridges 485 include a first passage portion or main body 493and a second passage portion or tab 492 transverse to the main body 493,such as a distal end 491 of the main body 493 (shown in FIG. 5B, withtwo ridges 485 shown in FIG. 5C and seal member 482 omitted forillustrative purposes). In alternative embodiments, each ridge 485 has amain body 493 and a tab 492. As shown in FIGS. 5A and 5B, the tab 492can extend outboard of the seal member 482 in a radial direction R. Thetabs 492 can be radially aligned with the main body 493, as depicted bytab 492′ and main body 493′, or axially offset or misaligned as depictedby tab 492″ and main body 493″ in FIG. 5A depending on a desiredpressure drop through corresponding cooling passages 487 and heataugmentation characteristics. The tabs 492 are configured to selectivelyengage a seal member 482 to bound or limit movement of the seal member482 relative to the ridges 485 in a circumferential or thicknessdirection T.

As shown in FIG. 5B, the seal member 482 can be dimensioned to be spacedfrom tabs 492A, 492B such that a length of cooling passages 487 isreduced. In alternative embodiments, the seal member 482 is configuredto abut both sets of tabs 492A, 492B. One or more radial retentionfeatures 499 can extend from corresponding tabs 492 (one radialretention feature 499 shown in FIG. 5C for illustrative purposes), suchas in a circumferential direction relative to the tabs 492. The tabs 492are configured to abut an outer surface 497 of the seal member 482 toretain the seal member 482 within a desired radial position, as shown bytabs 492A, 492B in FIG. 5B, such that a likelihood of the seal member482 becoming unseated during operation of the engine 20 is reduced. Inthe illustrative example, the radial retention features 499 can be abuta first or outer surface 497 of the seal member 482 opposite from asecond or inner surface 477 of the seal member 482 abutting the firstpassage portion or main body 493. The number of radial retentionfeatures 499 can be selected depending on a desired pressure dropthrough corresponding cooling passages 487 and heat augmentationcharacteristics to surrounding portions of the airfoil 461.

An arrangement similar to the cooling arrangement 478 for airfoil 461can be utilized for a vane, for example, as illustrated by coolingarrangement 778 of FIGS. 8A and 8B. Adjacent vanes 770A, 770B eachinclude a slot 794A, 794B extending from mate face 780A, 780B. Each slot794A, 794 is configured to receive a portion of seal member 782. One ormore ridges 785 include a tab 792 extending in a radial direction R froma main body 793. The tabs 792 space the seal member 782 from walls ofeach slot 794 in a circumferential or thickness direction T to limitrelative movement of the seal member 782. A cooling arrangement similarto cooling arrangement 778 of FIGS. 8A and 8B can be utilized for one ormore BOAS 869 or combustor panels 863, illustrated by a coolingarrangement 878 of FIG. 9.

Referring to FIG. 9A, the mate faces 780 can be continuous alongportions of the leakage gap 881, as illustrated by mate face 880A.Portions of the mate faces 780 can be discontinuous or segmented alongthe leakage gap 881 to define one or more flow passages 851, asillustrated by mate face 880B. The flow passages 851 provide multiplecooling flow sources or source pressures to the cooling passages 887rather than a single source pressure provided by adjacent mate faceshaving a continuous arrangement. The flow passages 851 reduce an overallpressure drop through the cooling arrangement 878, thereby reducing aquantity of cooling airflow corresponding to a desired heat augmentationand also reducing a likelihood of entrainment of hot combustion productsfrom the core flow path C entering into the leakage gap 881.

Although particular step sequences are shown, described, and claimed, itshould be understood that steps may be performed in any order, separatedor combined unless otherwise indicated and will still benefit from thepresent disclosure.

It should be understood that relative positional terms such as“forward,” “aft,” “upper,” “lower,” “above,” “below,” and the like arewith reference to the normal operational attitude of the vehicle andshould not be considered otherwise limiting.

The foregoing description is exemplary rather than defined by thelimitations within. Various non-limiting embodiments are disclosedherein, however, one of ordinary skill in the art would recognize thatvarious modifications and variations in light of the above teachingswill fall within the scope of the appended claims. It is therefore to beunderstood that within the scope of the appended claims, the disclosuremay be practiced other than as specifically described. For that reasonthe appended claims should be studied to determine true scope andcontent.

What is claimed is:
 1. A component for a gas turbine engine, comprising:a body including a cold side surface adjacent to a mate face; aplurality of ridges extending from the cold side surface; and a sealmember abutting the plurality of ridges to define a plurality of coolingpassages, the seal member moveable between a first position and a secondposition relative to the plurality of ridges when in an installedposition, each of the plurality of cooling passages including a firstinlet defined at the first position and a second, different inletdefined at the second position, and each of the plurality of coolingpassages includes an outlet adjacent to the mate face; wherein each ofthe plurality of ridges includes an elongated main body and a retentiontab, the main body extends in a circumferential direction along the coldside surface between opposed first and second ends, the respectiveoutlet is established at the second end, and the retention tab extendsfrom the first end such that the retention tab extends outward of theseal member in a radial direction; and wherein the retention tabs of theplurality of ridges are spaced apart from one another in an axialdirection and are arranged in a row to bound movement of the seal memberalong the plurality of ridges in the circumferential direction.
 2. Thecomponent as recited in claim 1, wherein the seal member is dimensionedsuch that the seal member is spaced a distance from the first end whenthe seal member is in the first position.
 3. The component as recited inclaim 1, wherein the plurality of ridges each include a radial retentionfeature extending in the circumferential direction from a respective oneof the retention tabs towards the second end such that the seal memberis trapped in the radial direction between the main body and the radialretention feature.
 4. The component as recited in claim 3, wherein: eachof the plurality of ridges has a C-shaped geometry established by themain body, the retention tab and the radial retention feature; andwherein the radial retention features of the plurality of ridges arespaced apart from one another in the axial direction.
 5. The componentas recited in claim 3, wherein the radial retention feature of at leastone of the retention tabs is misaligned in the axial direction from therespective main body.
 6. The component as recited in claim 3, whereinthe seal member is a feather seal dimensioned to extend outwardly of theplurality of ridges with respect to the axial direction.
 7. Thecomponent as recited in claim 1, wherein the cold side surface islocated at a slot extending inwardly from the mate face, and the slotreceives the seal member.
 8. The component as recited in claim 1,wherein the component is one of an airfoil, a blade outer air seal(BOAS), and a combustor panel.
 9. The component as recited in claim 8,wherein the airfoil includes an airfoil section extending in the radialdirection from a platform, and the cold side surface is located at anundersurface of the platform.
 10. The component as recited in claim 9,wherein the airfoil is a turbine blade.
 11. A gas turbine engine,comprising: a first component including a first set of ridges protrudingfrom a first cold side surface adjacent to a first mate face; a secondcomponent including a second set of ridges protruding from a second coldside surface adjacent to a second mate face, the second mate facecircumferentially adjacent to the first mate face relative to acircumferential direction to define a leakage gap; and a feather sealabutting the first set of ridges to define a first set of coolingchannels and abutting the second set of ridges to define a second set ofcooling channels such that the feather seal spaced apart from the firstcold side surface and the second cold side surface; wherein each of thefirst and second sets of ridges includes an elongated main body and aretention tab, the main body extends in the circumferential directionalong a respective one of the first and second cold side surfacesbetween opposed first and second ends, the second end is adjacent to arespective one of the first and second mate faces, and the retention tabextends in a radial direction from the first end such that the retentiontab extends outward of the feather seal; and wherein the retention tabsof the first set of ridges are spaced apart from one another in an axialdirection and are arranged in a row to bound movement of the featherseal along the first set of ridges in the circumferential direction. 12.The gas turbine engine as recited in claim 11, wherein each of the firstand second sets of cooling channels includes an inlet spaced apart fromthe leakage gap and an outlet adjacent to the leakage gap.
 13. The gasturbine engine recited in claim 11, wherein the first set of ridges eachinclude a radial retention feature extending in the circumferentialdirection from a respective one of the retention tabs towards therespective second end such that the seal member is trapped in the radialdirection between the main body and the radial retention feature. 14.The gas turbine engine as recited in claim 11, wherein the feather sealdefines a first width in the circumferential direction, and the firstends of the first and seconds set of ridges define a second width in thecircumferential direction, and a ratio of the first width to the secondwidth is equal to or less than 0.8.
 15. The gas turbine engine asrecited in claim 11, wherein each of the first component and the secondcomponent is one of an airfoil and a blade outer air seal (BOAS). 16.The gas turbine engine as recited in claim 14, wherein the firstcomponent is an airfoil, the airfoil includes an airfoil sectionextending in the radial direction from a platform, the platform includesan upper surface bounding a core flow path and an undersurface boundinga cooling cavity, and the first cold side surface is located at theundersurface of the platform.
 17. The gas turbine engine as recited inclaim 16, wherein the main body is raised from the undersurface of theplatform such that the feather seal is spaced apart from theundersurface relative to the radial direction.
 18. A method of sealingbetween adjacent components of a gas turbine engine, comprising:positioning a feather seal across a leakage gap defined between matefaces of adjacent components, and along a plurality of ridges to definea plurality of cooling passages; wherein the plurality of ridgesprotrude from cold side surfaces of the adjacent components such thatthe feather seal is spaced apart from the cold side surfaces; whereineach of the plurality of ridges includes an elongated main body and aretention tab, the main body extends in a circumferential directionalong the cold side surface between opposed first and second ends, arespective outlet is established at the second end, and the retentiontab extends in a radial direction from the first end such that theretention tab extends outward of the feather seal; wherein the retentiontabs of the plurality of ridges are spaced apart from one another in anaxial direction and are arranged in a row to bound movement of thefeather seal along the plurality of ridges in the circumferentialdirection; wherein each of the plurality of cooling passages includes aninlet that is spaced apart from the mate faces at each position of thefeather seal relative to the plurality of ridges; and moving the featherseal in the circumferential direction between a first circumferentialposition and a second circumferential position relative to the pluralityof ridges when in an installed position, and wherein the plurality ofridges are arranged such that the inlet is established at the firstcircumferential position and the inlet is established at the secondcircumferential position.
 19. The method as recited in claim 18,comprising communicating coolant from a cooling cavity to the pluralityof cooling passages when an edge face of the feather seal is alignedwith the first end of at least some of the plurality of ridges.
 20. Themethod as recited in claim 18, wherein the plurality of ridges eachinclude a radial retention feature extending in the circumferentialdirection from a respective one of the retention tabs towards therespective second end such that the seal member is trapped in the radialdirection between the main body and the radial retention feature.